Analysis and Comparison of Effects of an Airfoil or a Rod on Supersonic Cavity Flow.

The effects of an airfoil at different angles of attack and a circular cylindrical rod within the edge of the boundary layer flow at the leading edge of a cavity as a device for controlling the large pressure fluctuations (resonance tones) in the cavity were investigated. The airfoil results were compared with the rod in crossflow method positioned at the same leading edge location. The cavity used for testing corresponded to a length to depth ratio, L/D of 11.0/2.25 with a length to width ratio, L/W of 11.0/3.00 at a freestream Mach 1.84 flow. The study included measurements of dynamic pressure transducer output at 40 kHz and Frequency Spectra calculations, using Schlieren techniques for shock wave structures with velocity and vorticity fields obtained from PIV measurements. All airfoil configurations experienced flow separation to varying degrees. The negative 10 degree angle of attack configuration experienced the greatest amount of flow separation. All airfoil configurations provided varying degrees of cavity (resonant) tone suppression. Of the airfoil configurations, the negative 10 degree airfoil provided the best noise suppression with a 5 dB SPL reduction in broadband noise and a 9 dB reduction in peak amplitude for the 3rd resonant mode. Although all the airfoil configurations provided various levels of noise suppression, none of the configurations performed to the level of the rod in crossflow technique which provided an 8 dB SPL reduction in broadband noise and a 22 dB reduction in peak amplitude for the 2nd resonant mode. Indications of shear flow lofting effects could not be studied within any of the configurations tested. Lofting effect testing would have required flow field visualization of the cavity trailing edge region. Dynamic pressure measurements at a location near the cavity trailing edge did not detect the rod vortex shedding frequency, clearly. Because PIV results showed strong indication of vortex shedding, the lack of vortex shedding frequency data was attributed to the dynamic pressure transducer being located a far distance of 44 rod diameters downstream of the rod location. All airfoil test configurations showed evidence of deflections to the cavity leading edge oblique shock wave. The mechanisms of the deflection were the airfoil trailing edge shocks interacting with the cavity leading edge shock.